US4541775A - Clearance control in turbine seals - Google Patents

Clearance control in turbine seals Download PDF

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Publication number
US4541775A
US4541775A US06/480,519 US48051983A US4541775A US 4541775 A US4541775 A US 4541775A US 48051983 A US48051983 A US 48051983A US 4541775 A US4541775 A US 4541775A
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United States
Prior art keywords
sealing
rotor
disks
vanes
cooperating
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
US06/480,519
Inventor
Edward J. Hovan
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Raytheon Technologies Corp
Original Assignee
United Technologies Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by United Technologies Corp filed Critical United Technologies Corp
Priority to US06/480,519 priority Critical patent/US4541775A/en
Assigned to UNITED TECHNOLOGIES CORPORATION, A CORP OF DE. reassignment UNITED TECHNOLOGIES CORPORATION, A CORP OF DE. ASSIGNMENT OF ASSIGNORS INTEREST. Assignors: HOVAN, EDWARD J.
Priority to GB08402943A priority patent/GB2137283B/en
Priority to FR8402565A priority patent/FR2543616B1/en
Priority to JP59035861A priority patent/JPH0646003B2/en
Priority to DE3407218A priority patent/DE3407218C2/en
Application granted granted Critical
Publication of US4541775A publication Critical patent/US4541775A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
    • F01D11/025—Seal clearance control; Floating assembly; Adaptation means to differential thermal dilatations
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
    • F01D11/16—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing by self-adjusting means
    • F01D11/18—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing by self-adjusting means using stator or rotor components with predetermined thermal response, e.g. selective insulation, thermal inertia, differential expansion
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades
    • F01D5/084—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades the fluid circulating at the periphery of a multistage rotor, e.g. of drum type
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12—Cooling of plants
    • F02C7/16—Cooling of plants characterised by cooling medium
    • F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the invention is concerned with controlling the clearance in the tip and interstage seals in a multistage gas turbine.
  • a feature of the invention is a seal construction in which the clearance is controlled by thermal changes in the vane length resulting from engine operating conditions.
  • Another feature is a case supported seal cooperating with a rotor seal to control seal clearance.
  • the multistage rotor has its inner portion enclosed in a chamber formed by a sleeve within the rotor and sealing sleeves between the disks of the rotor.
  • This sleeve within the rotor also forms a pressurized compartment radially inward of the sleeve. Cooling fluid from the compartment enters the chamber by passing a seal which is made up of an outer sealing surface carried by the inner ends of the row of stator vanes and an inwardly extending sealing flange carried by the rotor.
  • a seal which is made up of an outer sealing surface carried by the inner ends of the row of stator vanes and an inwardly extending sealing flange carried by the rotor.
  • the seal clearance decreases in order to decrease the flow of cooling fluid to the rotor.
  • the single figure is a fragmentary sectional view through a turbine embodying the invention.
  • the case 2 for the turbine supports several rows of vanes 4, 6 and 8, with row 4 being the inlet stage of the multistage turbine.
  • Alternating with the rows of vanes are rows of blades 12, 14 and 16 carried by annular disks 18, 20 and 22 interconnected to one another to form the rotor by cooperating flanges 24 on opposite sides of each of the disks.
  • a flange 26 on the upstream side of the first rotor disk 18 is secured to and supports the upstream end of a sleeve 28 extending through the rotor and secured at its other end to the rotor cone 30.
  • the sleeve 28 forms between it and the rotor flanges 24 a chamber 31 through which cooling fluid flows for cooling the rotor.
  • Radially inward of the sleeve 28 is a compartment 32 which is pressurized with cooling fluid for use in cooling the turbine.
  • sealing sleeves 34 which carry on their outer surface sealing ribs 36 cooperating with a sealing sleeve 38 carried by the inner ends of the rows of stator vanes 6 and 8. Cooling air from the chamber 31 passes the flanges 24 through openings 40 therein and through openings 42 in the sealing sleeves to be directed over the surface of the rotor disk by a flange 44 carried by the sealing sleeve. These sealing sleeves are supported between the discs on flanges 45 on the adjacent surfaces of the discs.
  • the row of inlet vanes 4 carries a sealing ring 46 at the inner ends of the vanes to cooperate with a sealing member 48 carried by the flange 26 and the cooperating sleeve 28.
  • the sealing member 48 is H shape in cross section and has axially spaced annular disks 50 carrying on their outer peripheries seal elements 52 cooperating with an internal cylindrical sealing member 54 on the sealing element 46.
  • the sealing member 48 also carries on the inner peripheries of the disks 50 sealing fins 56 cooperating with an outwardly facing sealing ring 58.
  • the space between the two sealing disks communicates with the chamber 31 through a plurality of openings 60 in one of the disks 50 and overlapping flanges 61 and 62 on the cooperating flange 26 and sleeve 28.
  • the clearance between the inner seal 58 will be increased as the turbine vanes are heated and the space adjacent the sealing surface 54 will be decreased at the same time. Further as the turbine power is cut back and the vanes cool down in temperature as, for example, in power reduction to cruise, the clearance at the sleeve 58 will be decreased and the clearance at the sleeve 54 will be increased. A reduction in the clearance at sleeve 58 will reduce the flow of cooling fluid to the rotor chamber 31.
  • the spacing of the seals at the inner ends of the disks 50 thus serves as a controlled passage for the flow of cooling fluid to the rotor and the cooling fluid from compartment 32 passes the downstream seal disk 50, flows between the disks and through the openings 60 into the chamber 31 so that it flows around the inner portions of the rotor disks and through the openings 40 and 42 to be directed over the surface of the rotor disks for cooling.
  • the effect of the device is that as the turbine temperature is reduced by a cutback in power as, for example, from full power to cruise, the clearance at the inner ends of the sealing disk will be reduced and the flow of cooling fluid into the chamber 31 will be reduced with the result that a cooling of the turbine disks is reduced. Since the turbine disks are, therefore, not cooled as much, the clearance existing between the seal fins 36 and the cooperating sealing sleeve 38 in the interstage seals will be kept nearly the same or may, in fact, even be reduced for minimizing the escape of cooling fluid at this point.
  • the tip seals 63 located at the tips of the turbine blades 12 will be kept closer to the surrounding case structure 65, thereby tightening the tip seal clearance and reducing the air loss past these seals. This will further improve the low power efficiency.
  • the face of the forward sealing disk 50 and the face of the sealing ring 46 may have aligned small diameter openings 64 and 66 therein for permitting a minimum flow of cooling fluid regardless of the clearance at the inner ends of the sealing disks.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

A multistage turbine construction in which the cooling flow to the turbine rotor is controlled by cooperating sealing elements carried respectively by the row of vanes and by the rotor with the sealing surface carried by the vanes, the stationary surface being an outwardly sealing surface and the fins cooperating therewith carried by the rotor extending radially inward such that as the turbine rotor cools down during reduced power operation the sealing effect will be increased to reduce the flow of cooling fluid over the rotor surface.

Description

DESCRIPTION
1. Technical Field
The invention is concerned with controlling the clearance in the tip and interstage seals in a multistage gas turbine.
2. Background Art
Maintaining seal clearance over the entire flight spectrum is difficult. The thermal shrinkage of the airfoils as power is pulled back at cruise, for example, leads to large clearances and inefficiency of the turbine. This is particularly true where the tip seals and inner supports are tied to the case as opposed to a floating mount. It is desirable to maintain a minimum clearance at all conditions of operation or to maintain such a clearance as to minimize any losses resulting from excessive seal clearances.
DISCLOSURE OF INVENTION
A feature of the invention is a seal construction in which the clearance is controlled by thermal changes in the vane length resulting from engine operating conditions. Another feature is a case supported seal cooperating with a rotor seal to control seal clearance.
According to the invention the multistage rotor has its inner portion enclosed in a chamber formed by a sleeve within the rotor and sealing sleeves between the disks of the rotor. This sleeve within the rotor also forms a pressurized compartment radially inward of the sleeve. Cooling fluid from the compartment enters the chamber by passing a seal which is made up of an outer sealing surface carried by the inner ends of the row of stator vanes and an inwardly extending sealing flange carried by the rotor. As the seal clearance varies by changes in the vane temperature the amount of cooling fluid is varied to suit the turbine operation. Specifically, as the vanes cool and contract when power is reduced from max power to cruise, for example, the seal clearance decreases in order to decrease the flow of cooling fluid to the rotor. By so doing the cooling effect on the disks of the rotor is reduced and this results in a tightening of the interstage rotor seals positioned between the rows of vanes and the rotor structure.
The foregoing and other objects, features and advantages of the present invention will become more apparent in the light of the following detailed description of preferred embodiments thereof as shown in the accompanying drawing.
BRIEF DESCRIPTION OF THE DRAWING
The single figure is a fragmentary sectional view through a turbine embodying the invention.
BEST MODE FOR CARRYING OUT THE INVENTION
The case 2 for the turbine supports several rows of vanes 4, 6 and 8, with row 4 being the inlet stage of the multistage turbine.
Alternating with the rows of vanes are rows of blades 12, 14 and 16 carried by annular disks 18, 20 and 22 interconnected to one another to form the rotor by cooperating flanges 24 on opposite sides of each of the disks. A flange 26 on the upstream side of the first rotor disk 18 is secured to and supports the upstream end of a sleeve 28 extending through the rotor and secured at its other end to the rotor cone 30. The sleeve 28 forms between it and the rotor flanges 24 a chamber 31 through which cooling fluid flows for cooling the rotor. Radially inward of the sleeve 28 is a compartment 32 which is pressurized with cooling fluid for use in cooling the turbine. Extending between adjacent rotor disks radially outward from the flanges 24 are sealing sleeves 34 which carry on their outer surface sealing ribs 36 cooperating with a sealing sleeve 38 carried by the inner ends of the rows of stator vanes 6 and 8. Cooling air from the chamber 31 passes the flanges 24 through openings 40 therein and through openings 42 in the sealing sleeves to be directed over the surface of the rotor disk by a flange 44 carried by the sealing sleeve. These sealing sleeves are supported between the discs on flanges 45 on the adjacent surfaces of the discs.
The row of inlet vanes 4 carries a sealing ring 46 at the inner ends of the vanes to cooperate with a sealing member 48 carried by the flange 26 and the cooperating sleeve 28. The sealing member 48 is H shape in cross section and has axially spaced annular disks 50 carrying on their outer peripheries seal elements 52 cooperating with an internal cylindrical sealing member 54 on the sealing element 46.
The sealing member 48 also carries on the inner peripheries of the disks 50 sealing fins 56 cooperating with an outwardly facing sealing ring 58. The space between the two sealing disks communicates with the chamber 31 through a plurality of openings 60 in one of the disks 50 and overlapping flanges 61 and 62 on the cooperating flange 26 and sleeve 28.
With the sealing member 48 positioned as it is between the opposed sealing sleeves 54 and 58, it will be apparent that the clearance between the inner seal 58 will be increased as the turbine vanes are heated and the space adjacent the sealing surface 54 will be decreased at the same time. Further as the turbine power is cut back and the vanes cool down in temperature as, for example, in power reduction to cruise, the clearance at the sleeve 58 will be decreased and the clearance at the sleeve 54 will be increased. A reduction in the clearance at sleeve 58 will reduce the flow of cooling fluid to the rotor chamber 31.
The spacing of the seals at the inner ends of the disks 50 thus serves as a controlled passage for the flow of cooling fluid to the rotor and the cooling fluid from compartment 32 passes the downstream seal disk 50, flows between the disks and through the openings 60 into the chamber 31 so that it flows around the inner portions of the rotor disks and through the openings 40 and 42 to be directed over the surface of the rotor disks for cooling.
Essentially the effect of the device is that as the turbine temperature is reduced by a cutback in power as, for example, from full power to cruise, the clearance at the inner ends of the sealing disk will be reduced and the flow of cooling fluid into the chamber 31 will be reduced with the result that a cooling of the turbine disks is reduced. Since the turbine disks are, therefore, not cooled as much, the clearance existing between the seal fins 36 and the cooperating sealing sleeve 38 in the interstage seals will be kept nearly the same or may, in fact, even be reduced for minimizing the escape of cooling fluid at this point. With a decreased cooling of the turbine disks the tip seals 63 located at the tips of the turbine blades 12 will be kept closer to the surrounding case structure 65, thereby tightening the tip seal clearance and reducing the air loss past these seals. This will further improve the low power efficiency.
The face of the forward sealing disk 50 and the face of the sealing ring 46 may have aligned small diameter openings 64 and 66 therein for permitting a minimum flow of cooling fluid regardless of the clearance at the inner ends of the sealing disks.
Although the invention has been shown and described with respect to a preferred embodiment thereof, it should be understood by those skilled in the art that other changes and omissions in the form and detail thereof may be made therein without departing from the spirit and scope of the invention.

Claims (4)

I claim:
1. In a gas turbine:
a rotor having at least two disks;
seal sleeves extending between the disks and supported by flanges on the discs and having sealing elements on their outer periphery;
one end disk carrying a sealing member having at least one inwardly extending sealing flange;
a row of vanes adjacent one end disk and at least one other row of vanes between the disks;
a sealing ring carried by the inner ends of said row of vanes and having an outer sealing surface cooperating with said sealing flange;
a sleeve extending through the rotor from said sealing member and forming with the sealing member and sealing sleeves a chamber enclosing the inner portions of the rotor disks;
said sleeve also forming inwardly of the sleeve a pressurized compartment for cooling fluid for the rotor, said cooperating sealing flange and outer sealing surface providing a clearance space for the flow of fluid from the compartment to enter the chamber.
2. A gas turbine as in claim 1 in which the sealing member has a plurality of passages therethrough for the flow of cooling fluid into the chamber.
3. A gas turbine as in claim 1 in which the sealing member has a pair of spaced sealing flanges cooperating with the outer sealing surface and the space between the sealing flanges communicates with the chamber.
4. A gas turbine as in claim 1 in which the rows of vanes between the disks carry sealing sleeves cooperating with fins on the rotor sleeves positioned between the disks.
US06/480,519 1983-03-30 1983-03-30 Clearance control in turbine seals Expired - Lifetime US4541775A (en)

Priority Applications (5)

Application Number Priority Date Filing Date Title
US06/480,519 US4541775A (en) 1983-03-30 1983-03-30 Clearance control in turbine seals
GB08402943A GB2137283B (en) 1983-03-30 1984-02-03 Clearance control in turbine seals
FR8402565A FR2543616B1 (en) 1983-03-30 1984-02-21 CHECKING THE FREE SPACE IN THE TURBINE GASKETS
JP59035861A JPH0646003B2 (en) 1983-03-30 1984-02-27 Gas turbine sealing structure
DE3407218A DE3407218C2 (en) 1983-03-30 1984-02-28 Gas turbine

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Application Number Priority Date Filing Date Title
US06/480,519 US4541775A (en) 1983-03-30 1983-03-30 Clearance control in turbine seals

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US4541775A true US4541775A (en) 1985-09-17

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US (1) US4541775A (en)
JP (1) JPH0646003B2 (en)
DE (1) DE3407218C2 (en)
FR (1) FR2543616B1 (en)
GB (1) GB2137283B (en)

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US4841726A (en) * 1985-11-19 1989-06-27 Mtu-Munchen Gmbh Gas turbine jet engine of multi-shaft double-flow construction
US5054996A (en) * 1990-07-27 1991-10-08 General Electric Company Thermal linear actuator for rotor air flow control in a gas turbine
US5232339A (en) * 1992-01-28 1993-08-03 General Electric Company Finned structural disk spacer arm
US5320484A (en) * 1992-08-26 1994-06-14 General Electric Company Turbomachine stator having a double skin casing including means for preventing gas flow longitudinally therethrough
US5779436A (en) * 1996-08-07 1998-07-14 Solar Turbines Incorporated Turbine blade clearance control system
EP1004759A2 (en) * 1998-11-24 2000-05-31 General Electric Company Bay cooled turbine casing
EP0924388A3 (en) * 1997-12-19 2000-08-16 Rolls-Royce Deutschland GmbH System to keep the blade tip clearance in a gas turbine constant
US6116852A (en) * 1997-12-11 2000-09-12 Pratt & Whitney Canada Corp. Turbine passive thermal valve for improved tip clearance control
US6638013B2 (en) 2002-02-25 2003-10-28 Honeywell International Inc. Thermally isolated housing in gas turbine engine
US6719524B2 (en) 2002-02-25 2004-04-13 Honeywell International Inc. Method of forming a thermally isolated gas turbine engine housing
US20070003410A1 (en) * 2005-06-23 2007-01-04 Siemens Westinghouse Power Corporation Turbine blade tip clearance control
US20120134779A1 (en) * 2010-11-29 2012-05-31 Alexander Anatolievich Khanin Gas turbine of the axial flow type
CN102985639A (en) * 2010-07-13 2013-03-20 西门子公司 Sealing element for sealing a gap
US20160102569A1 (en) * 2014-10-14 2016-04-14 Alstom Technology Ltd Steam turbine gland arrangement
US20170030196A1 (en) * 2015-07-28 2017-02-02 MTU Aero Engines AG Gas turbine
US11415062B2 (en) 2020-11-18 2022-08-16 Raytheon Technologies Corporation Rotating sleeve controlling clearance of seal assembly of gas turbine engine
GB2603884A (en) * 2004-09-07 2022-08-24 Rolls Royce Plc A cooling arrangement

Families Citing this family (11)

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Publication number Priority date Publication date Assignee Title
DE3428892A1 (en) * 1984-08-04 1986-02-13 MTU Motoren- und Turbinen-Union München GmbH, 8000 München Vane and sealing gap optimization device for compressors of gas turbine power plants, in particular gas turbine jet power plants
US4642024A (en) * 1984-12-05 1987-02-10 United Technologies Corporation Coolable stator assembly for a rotary machine
FR2600377B1 (en) * 1986-06-18 1988-09-02 Snecma DEVICE FOR MONITORING THE COOLING AIR FLOWS OF AN ENGINE TURBINE
US5271711A (en) * 1992-05-11 1993-12-21 General Electric Company Compressor bore cooling manifold
GB2307279B (en) * 1995-11-14 1999-11-17 Rolls Royce Plc A gas turbine engine
GB0722511D0 (en) 2007-11-19 2007-12-27 Rolls Royce Plc Turbine arrangement
FR2953251B1 (en) * 2009-12-01 2011-12-09 Snecma ROTOR DRUM COMPRISING A PLURALITY OF VENTILATION ORIFICES FOR IMPACT COOLING AND MOTOR COMPRISING SUCH A DRUM
FR2953250B1 (en) * 2009-12-01 2016-04-01 Snecma TURBINE ROTOR OF A TURBOMACHINE ITS MANUFACTURING PROCESS AND TURBOMACHINE EQUIPPED WITH SUCH A ROTOR
FR2954797B1 (en) * 2009-12-29 2016-03-18 Snecma LOW PRESSURE TURBINE ROTOR HAVING A REAR VENTILATION ARRANGEMENT TO THE FRONT OF AN ARROW DISC, AND TURBOMACHINE EQUIPPED WITH SUCH A ROTOR
US10001061B2 (en) 2014-06-06 2018-06-19 United Technologies Corporation Cooling system for gas turbine engines
EP3106613A1 (en) * 2015-06-06 2016-12-21 United Technologies Corporation Cooling system for gas turbine engines

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US2401826A (en) * 1941-11-21 1946-06-11 Dehavilland Aircraft Turbine
US2552239A (en) * 1946-10-29 1951-05-08 Gen Electric Turbine rotor cooling arrangement
GB738656A (en) * 1952-07-26 1955-10-19 Power Jets Res & Dev Ltd Blades for compressors, turbines and like bladed fluid flow machines
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US3291447A (en) * 1965-02-15 1966-12-13 Gen Electric Steam turbine rotor cooling
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US3989410A (en) * 1974-11-27 1976-11-02 General Electric Company Labyrinth seal system
US4005946A (en) * 1975-06-20 1977-02-01 United Technologies Corporation Method and apparatus for controlling stator thermal growth
US4184797A (en) * 1977-10-17 1980-01-22 General Electric Company Liquid-cooled turbine rotor
US4217755A (en) * 1978-12-04 1980-08-19 General Motors Corporation Cooling air control valve
US4242042A (en) * 1978-05-16 1980-12-30 United Technologies Corporation Temperature control of engine case for clearance control

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US4213738A (en) * 1978-02-21 1980-07-22 General Motors Corporation Cooling air control valve
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CA614162A (en) * 1961-02-07 S. Hall Russell Variable blade system
CA635902A (en) * 1962-02-06 General Electric Company Cooling means for a multi-stage turbine
US809816A (en) * 1904-07-29 1906-01-09 Gen Electric Elastic-fluid turbine.
US2401826A (en) * 1941-11-21 1946-06-11 Dehavilland Aircraft Turbine
US2552239A (en) * 1946-10-29 1951-05-08 Gen Electric Turbine rotor cooling arrangement
GB738656A (en) * 1952-07-26 1955-10-19 Power Jets Res & Dev Ltd Blades for compressors, turbines and like bladed fluid flow machines
US2787440A (en) * 1953-05-21 1957-04-02 Westinghouse Electric Corp Turbine apparatus
US2859934A (en) * 1953-07-29 1958-11-11 Havilland Engine Co Ltd Gas turbines
US2937495A (en) * 1956-02-27 1960-05-24 Power Jets Res & Dev Ltd Gas turbine plant
FR1207772A (en) * 1957-07-18 1960-02-18 Rolls Royce Improvements to fluid machines with paddle rotors
FR1237157A (en) * 1959-05-26 1960-07-29 Hispano Suiza Sa Improvements to compressor installations for gas streams to be partially recycled
US3043560A (en) * 1960-01-28 1962-07-10 Louis E Varadi Turbine cooling system
US3250512A (en) * 1962-11-09 1966-05-10 Rolls Royce Gas turbine engine
US3224194A (en) * 1963-06-26 1965-12-21 Curtiss Wright Corp Gas turbine engine
US3291447A (en) * 1965-02-15 1966-12-13 Gen Electric Steam turbine rotor cooling
US3527053A (en) * 1968-12-11 1970-09-08 Gen Electric Gas turbine engine with improved gas seal
US3826084A (en) * 1970-04-28 1974-07-30 United Aircraft Corp Turbine coolant flow system
US3959966A (en) * 1973-10-05 1976-06-01 Rolls-Royce (1971) Limited Gas turbine engine
US3989410A (en) * 1974-11-27 1976-11-02 General Electric Company Labyrinth seal system
US4005946A (en) * 1975-06-20 1977-02-01 United Technologies Corporation Method and apparatus for controlling stator thermal growth
US4184797A (en) * 1977-10-17 1980-01-22 General Electric Company Liquid-cooled turbine rotor
US4242042A (en) * 1978-05-16 1980-12-30 United Technologies Corporation Temperature control of engine case for clearance control
US4217755A (en) * 1978-12-04 1980-08-19 General Motors Corporation Cooling air control valve

Cited By (25)

* Cited by examiner, † Cited by third party
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GB2137283A (en) 1984-10-03
GB2137283B (en) 1987-03-04
DE3407218C2 (en) 1994-12-15
FR2543616B1 (en) 1986-09-12
JPH0646003B2 (en) 1994-06-15
JPS59180008A (en) 1984-10-12
DE3407218A1 (en) 1984-10-04
FR2543616A1 (en) 1984-10-05
GB8402943D0 (en) 1984-03-07

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